Entries |
Document | Title | Date |
20080202094 | AEROENGINE FITTED WITH HEAT EXCHANGER MEANS - The invention relates to an aeroengine comprising:
| 08-28-2008 |
20080236137 | ACOUSTIC FLOW STRAIGHTENER FOR TURBOJET ENGINE FAN CASING - The present invention relates to a structural or non-structural turbojet engine fan casing connecting arm ( | 10-02-2008 |
20080236138 | Arrangement for Propelling an Aircraft, Aircraft and Outlet Nozzle for a Jet Engine - The invention relates to an arrangement for propelling an aircraft comprising a jet engine and an outlet nozzle arranged downstream of the jet engine. The jet engine is of a type which generates an internal core flow and an external fan flow. A part of the outlet nozzle has an internal surface which defines a gas duct for the fan flow. The aforementioned part of the outlet nozzle includes a wall structure that is arranged at a distance from the internal surface such that it separates the gas duct for the fan flow from an internal gas duct for the core flow. | 10-02-2008 |
20080245053 | GAS TURBINE ENGINE DIFFUSER AND COMBUSTION CHAMBER AND GAS TURBINE ENGINE COMPRISING SAME - The present invention relates to a gas turbine engine diffuser ( | 10-09-2008 |
20080271431 | Turbojet Engine with Attenuated Jet Noise - According to the invention, a plurality of hatches ( | 11-06-2008 |
20080302083 | INTERNAL MIXING OF A PORTION OF FAN EXHAUST FLOW AND FULL CORE EXHAUST FLOW IN AIRCRAFT TURBOFAN ENGINES - An aircraft includes at least one turbofan engine assembly having a shrouded core engine, a short nacelle surrounding a fan and a forward portion of the core engine, and a fan exhaust duct through the nacelle. A mixer duct shell is positioned substantially coaxial with the engine shroud and extends forwardly into the fan duct to provide an interstitial mixer duct between the mixer duct shell and the core engine shroud. The aft portion of the mixer duct shell extends over a turbine exhaust frame, an attached mixer (if included), and a tail cone exhaust plug. The mixer duct shell can reduce noise and plume exhaust heat radiated from aircraft turbofan engines. | 12-11-2008 |
20090000272 | TANGENTIAL ANTI-SWIRL AIR SUPPLY - A turbofan flow delivery system includes a fan case housing a fan. Flow exit guide vanes are arranged downstream from the fan and extend radially inwardly from the fan case toward a bypass flow path. A supply passage includes an opening provided in the fan case between the fan and flow exit guide vanes configured to selectively provide pressurized air to a component using air from the bypass flow path. Swirling air from the bypass flow path enters the supply passage and is converted to a static pressure. | 01-01-2009 |
20090056306 | GAS TURBINE ENGINE FRONT ARCHITECTURE - A turbine engine is disclosed that includes a fan case surrounding a fan. A core is supported relative to the fan case by support structure, such as flow exit guide vanes, which are arranged downstream from the fan. The core includes a core housing having an inlet case arranged to receive airflow from the fan. A compressor case is arranged axially adjacent to the inlet case and surrounds a compressor stage. In one example, the example turbine engine includes a gear train arranged between the fan and a spool. The gear train is axially aligned and supported by the inlet case. An intermediate case is arranged axially adjacent to the compressor case. The support structure is arranged axially forward of the intermediate case. In one example, the support structure is axially aligned with the compressor case. | 03-05-2009 |
20090071121 | COUNTER-ROTATING GEARBOX FOR TIP TURBINE ENGINE - A tip turbine engine ( | 03-19-2009 |
20090077946 | Fan control apparatus - A fan control apparatus includes a fan, two engine+compressor combinations, two air supply systems, and an FCC. When an abnormality occurs in one of the air supply systems and one of the engine+compressor combinations, the FCC maintains the flow rate of the normally operating air supply system and then increases the flow rate. As a result, the normally operating drive source is prevented from overloading. In another embodiment, a fan control apparatus includes a fan, an air source, two air supply systems, and a bypass channel. The air is caused to flow through the bypass channel when an abnormality occurs in one of the air supply systems. As a result, the time that elapses till the fluid can be supplied at a necessary flow rate is shortened. | 03-26-2009 |
20090090095 | AIRCRAFT TURBOFAN ENGINE - According to the invention, there is arranged on the inner fan cowl ( | 04-09-2009 |
20090120058 | Tip Turbine Engine Integral Fan, Combustor, and Turbine Case - A tip turbine engine assembly includes an integral engine outer case located radially outward from a fan assembly. The integral outer case includes a rear portion and a forward portion with an arcuate portion that curves radially inwardly to form a compartment. An annular combustor is housed and mounted in the compartment. Fan inlet guide vanes are integrally formed with the arcuate portion to form the integral case portion. The rear portion, forward portion, and fan inlet guide vanes are integrally | 05-14-2009 |
20090145106 | TURBOFAN ENGINE UTILIZING AN AERODYNAMICALLY COUPLED PRE-COMBUSTION POWER TURBINE - A turbofan jet engine design that utilizes aerodynamic coupling to transmit power from a high-speed engine core to a lower speed fan, thereby simplifying the design and improving the thrust-to-weight ratio compared to previous turbofan designs. The engine uses a low speed co-rotating power turbine located upstream of the engine core to drive the fan. The high-speed core uses a centrifugal impeller to pressurize the inlet flow, the flow exits the high-speed impeller without diffusing its high-speed angular momentum and enters directly into the low speed co-rotating power turbine impeller. The incoming flow is then turned by the low speed co-rotating turbine impeller and exits the turbine opposing its direction of rotation, thereby extracting power from the flow to drive the fan. The exit flow from the low speed power turbine then enters the combustor inlet. This engine configuration enables the power turbine to be constructed from high strength, low-density materials (that are not suitable for use in higher temperature power turbines) thereby reducing the weight of the engine. This engine configuration also eliminates the need for the core compressor diffuser, the power turbine nozzle guide varies, and the low speed shaft that connects the fan to the power turbine in conventional turbofans. | 06-11-2009 |
20090178386 | Aircraft Propulsion System - The invention is a novel aircraft propulsion system architecture that delivers thrust at minimum fuel consumption rates with the side benefits of noise and pollution emission abatement. The invention implements three technologies:
| 07-16-2009 |
20090211221 | Auxiliary propulsor for a variable cycle gas turbine engine - A core engine of a variable cycle gas turbine engine includes a low pressure spool for generating streams of bypass air and pressurized air, and a high pressure spool for further pressurizing the stream of pressurized air to generate streams of combustion air and supercharged auxiliary air. A peripheral case surrounds the engine case to form a peripheral duct. An auxiliary combustor and propulsor are positioned within the peripheral duct. A bleed duct extends from the high pressure spool to the auxiliary combustor. Variable ductwork directs airflow through the bleed duct and peripheral duct in two modes. A first mode comprises directing the stream of auxiliary air to the auxiliary combustor, and directing stream of inlet air through the peripheral duct. A second mode comprises directing the stream of auxiliary air into the stream of bypass air, and preventing inlet air from entering the peripheral duct. | 08-27-2009 |
20090211222 | Rear propulsor for a variable cycle gas turbine engine - A gas turbine propulsion system comprises a turbofan engine, a peripheral duct, an annular frame, an auxiliary turbine and an auxiliary fan. The turbofan engine is configured to produce bypass air and combustion air. The bypass air flows through a bypass duct and the combustion air flows through a core engine. The peripheral duct surrounds the turbofan engine and is configured to selectively receive peripheral inlet air. The annular frame is disposed aft of the bypass duct and the peripheral duct, and is rotatable to alternately guide the bypass air out the bypass duct or the peripheral duct. The auxiliary turbine is connected to an aft end of the core engine and is configured to receive the combustion air. The auxiliary fan is connected to the auxiliary turbine and is configured to receive airflow from the peripheral duct. | 08-27-2009 |
20090288388 | Gas turbine exhaust - An exhaust device is disclosed that includes a duct that turns an exhaust flow from a gas turbine engine from a first direction to a second direction. A diverter cup is disposed within the duct and serves to split the exhaust flow into two streams as well as introduce cooling air through an ejector pump at the downstream end of the diverter cup. A splitter is disposed downstream of the diverter cup and serves to split the cooling air into two streams which are thereafter mixed with the split exhaust flow. A diffuser is created in the exhaust device between the duct and the splitter. Various cooling slots are also provided in the duct. | 11-26-2009 |
20090293449 | Gas-Turbine engine in particular aircraft engine - On a gas-turbine engine, in particular an aircraft engine, with a fan, a fan casing and a fan duct as well as with high-pressure and low-pressure turbines arranged behind each other in flow direction in the casing of the engine, the auxiliaries connected are to be operated with only a small increase in engine power. For this purpose in the flow direction of the airflow ( | 12-03-2009 |
20100000199 | MANAGING LOW PRESSURE TURBINE MAXIMUM SPEED IN A TURBOFAN ENGINE - (A2) A turbofan engine control system for managing a low pressure turbine speed is provided. The turbofan engine control system includes a low spool having a low pressure turbine that are housed in a core nacelle. The low pressure turbine is adapted to rotate at a speed and includes a maximum design speed. A turbofan is coupled to the low spool. A fan nacelle surrounds the turbofan and core nacelle and provides a bypass flow path. The bypass flow path includes a nozzle exit area. A controller is programmed to command a flow control device adapted to effectively decrease the nozzle exit area in response to a condition. Reducing the nozzle exit area, either physically or otherwise, maintains the speed below the maximum design speed. | 01-07-2010 |
20100011741 | GAS TURBINE ENGINE WITH DUAL FANS DRIVEN ABOUT A CENTRAL CORE AXIS - A gas turbine engine includes a compressor section, a combustor section and a turbine section, all arranged about a central axis. A turbine rotor drives a driveshaft for rotation on the central axis. The driveshaft in turn drives at least two fan rotors on axes parallel to, but spaced from, the central axis. The fan rotors each deliver a portion of the air they move into a central inlet for delivering the air to the compressor section. In addition, the majority of the air moved by the plurality of fan rotors passes between a cowl and an outer periphery of a housing for the central core of the engine. This bypass air provides propulsion as is known in the turbojet art. | 01-21-2010 |
20100031627 | Heater Assemblies, Gas Turbine Engine Systems Involving Such Heater Assemblies and Methods for Manufacturing Such Heater Assemblies - Heater assemblies, gas turbine engine systems involving such heater assemblies and methods for manufacturing such heater assemblies are provided. In this regard, a representative method includes: providing a substrate; forming a heating element using a thermal sprayed metal process such that the metal heater comprises an aggregation of sprayed metal particles; attaching the heating element to the substrate; and consolidating at least some of the metal particles located at an exterior of the heating element to form an electrical contact pad of the heating element. | 02-11-2010 |
20100031628 | JET FLOW DISCHARGE NOZZLE AND JET ENGINE - A jet flow discharge nozzle includes a tubular partition wall, a tubular casing which covers an outer periphery of the tubular partition wall, and a bifurcation which extends in an axial direction of the tubular partition wall and the casing so as to support downstream ends of the tubular partition wall and the casing from outside. Inside of the tubular partition wall is defined as a flow path for a high-speed core flow. A space between the tubular partition wall and the casing is defined as a flow path for a low-speed bypass flow. A pair of first projections is provided at a downstream edge of the tubular partition wall near the bifurcation at positions symmetric about the bifurcation. One of ridges of each of the first projections faces an upstream side. The invention reduces noise by avoiding adverse effects of the presence of the bifurcation on an increase in noise. | 02-11-2010 |
20100031629 | LOCKABLE GUIDING SYSTEM FOR A MOBILE PART OF A NACELLE - The invention relates to a guiding device ( | 02-11-2010 |
20100107597 | Method for Reducing Sound Output at the Back of a Turbo Engine and Turbo Engine Improved By This Method - The invention relates to a method for reducing the noise missions from the rear of a turbo engine, and a turbo engine improved by said method. According to the invention, the nozzle for the cold flow ( | 05-06-2010 |
20100139241 | FLOW DIVIDER FOR A FAN ENGINE - A flow divider ( | 06-10-2010 |
20100162682 | INLET PARTICLE SEPARATOR SYSTEM FOR A GAS TURBINE ENGINE - Embodiments include an inlet particle separator system for a gas turbine engine. The inlet particle separator system includes an inertial particle separator that separates incoming air into a cleaned air flow and a scavenge flow. Embodiments may also include an ejector that provides a draw on a scavenge duct and entrains the scavenge flow into a charged flow, e.g., such as the output of a first stage fan. The ejector may have a variable output. | 07-01-2010 |
20100180572 | FLADE FAN WITH DIFFERENT INNER AND OUTER AIRFOIL STAGGER ANGLES AT A SHROUD THEREBETWEEN - A FLADE fan assembly includes radially inner and outer airfoils extending radially inwardly and outwardly respectively from an annular shroud circumferentially disposed about a centerline. Inner and outer chords extend between inner and outer leading and trailing edges of inner and outer airfoil cross-sections of the radially inner and outer airfoils respectively. Inner and outer stagger angles between the inner and outer chords respectively at the shroud and the centerline are different. The radially outer airfoils may outnumber the radially inner airfoils and particularly by a ratio in a range of 1.5:1 to about 4:1. Load paths or radii may extend radially through the inner and outer airfoils and through the rotating shroud between the inner and outer airfoils and may pass near or through the inner and outer leading edges and through the inner and outer trailing edges. | 07-22-2010 |
20100186369 | DUAL FLOW TURBOSHAFT ENGINE AND IMPROVED HOT FLOW NOZZLE - According to the invention, the hot flow nozzle includes a single skin ( | 07-29-2010 |
20100192540 | CLOSING SYSTEM FOR AN AIRCRAFT FAN SHROUD - A closing system for a fan cowling on a jet engine of an aircraft that has at least one assembly including a male contrivance integral with the cowling, and a female contrivance integral with the jet engine that has a lower element forming at least one first guide ramp and an upper element forming at least one second guide ramp, with the aforementioned female contrivance being suitable on the one hand for guiding the male contrivance during the opening and/or the closing of the fan cowling, and on the other hand providing a limitation of the radial and circumferential displacements of the male contrivance. | 08-05-2010 |
20100205930 | EXHAUST SYSTEM FOR GAS TURBINE - The invention relates to an exhaust system ( | 08-19-2010 |
20100212285 | TURBOPROP PROPULSION UNIT WITH PUSHER PROPELLER - A turboprop propulsion unit includes at least one pusher propeller | 08-26-2010 |
20100223905 | SCOOP OF A RUNNING-GAP CONTROL SYSTEM OF AN AIRCRAFT GAS TURBINE - A scoop for a fairing | 09-09-2010 |
20100229526 | ATTACHEMENT OF A JET ENGINE NACELLE STRUCTURE BY MEANS OF A REINFORCED KNIFE-EDGE/GROOVE COUPLING - The present invention relates to a nacelle ( | 09-16-2010 |
20100251692 | Methods of combining a series of more efficient aircraft engines into a unit, or modular units - The present invention generally relates to units of engines and more particularly to units containing a unique combined-cycle (combustion-detonation) “counter-rotation, anti-gyration, gyroscopic,” turbine fan-jet/free-piston engine configuration for induced air supercharging and boosting the performance of novel Ramjet engines or Ramjet engine configurations by improving internal air-stream dynamics. These dynamics are the result of co-operative air stream intermixing through convergent, supercharge-attenuated, inducted, compressed, tuned, pre-heated ambient air. Achieved through the varying of the geometric structural form and the utilization of unique engines and air induction and propulsion conformations, aided with supplemental air, fuel, oxygen and optiomal water and electrolyte charging. | 10-07-2010 |
20100293920 | PROPFAN ENGINE - A propfan engine has a power-producing core, a plurality of nozzles for exhausting exhaust gas from the core, and a row of propeller blades rotatable about an axis. The blades are positioned downstream of the exhaust nozzles. The exhaust nozzles are rotatable about said axis. The nozzles can be aligned with the gaps between successive propeller blades in order to reduce noise during operation of the engine. | 11-25-2010 |
20100326046 | VALVE SYSTEM FOR A GAS TURBINE ENGINE - A valve system intermediate a secondary flow path and a primary flow path to selectively communicate secondary airflow into the primary gas flow path and control airflow injected from a higher pressure plenum into a lower pressure flowpath. | 12-30-2010 |
20110011058 | TURBOFAN ENGINE - With a turbofan engine the inner sidewall ( | 01-20-2011 |
20110030336 | COANNULAR DUCTED FAN - Secondary air flow is provided for a ducted fan having an engine core driving a fan blisk. The fan blisk incorporates a set of thrust fan blades extending from an outer hub and a set of integral secondary flow blades extending intermediate an inner hub and the outer hub. A nacelle provides a first flow duct for the thrust fan blades and a secondary flow duct carries flow from the integral secondary flow blades. | 02-10-2011 |
20110030337 | WALL COOLING DEVICE - The invention relates to a wall-cooling device for cooling a wall ( | 02-10-2011 |
20110047960 | DUAL-FLOW TURBINE ENGINE FOR AIRCRAFT WITH LOW NOISE EMISSION - The invention relates to a dual-flow turbine engine for an aircraft with low noise emission, wherein the opening ( | 03-03-2011 |
20110083416 | PROPULSION ENGINE - A propulsion engine ( | 04-14-2011 |
20110083417 | COOLING SYSTEM - A cooling system for an open rotor gas turbine engine is provided. The system comprises an engine exhaust, and a row of open rotor propeller blades located rearwardly of the exhaust such that hot exhaust gases impinge on the propeller blades. Each blade has an internal cooling network through which cooling air flows to cool the blade in respect of the hot exhaust gases. Each blade also has one or more intakes which provide cooling air to the cooling network. The or each intake is located radially outwardly of the position of impingement of the hot exhaust gas on the blade. | 04-14-2011 |
20110094204 | COUPLING SYSTEM CONNECTING AN INTERNAL STRUCTURE AND AN EXTERNAL STUCTURE OF A JET ENGINE NACELLE - The present invention relates to a turbofan jet engine nacelle intended to be attached to a structure of an aircraft by an engine strut and comprising a forward air inlet section, a mid-section intended to surround a jet engine fan, and an aft section having an internal structure intended to serve as a casing for an aft portion of the jet engine and, together with an external structure ( | 04-28-2011 |
20110138772 | TURBINE DETUNER FOR RECOVERING KINETIC ENERGY FROM GAS TURBINE ENGINE EXHAUST GASES - A gas turbine engine test cell has a turbine detuner capable of recovering kinetic energy from exhaust gases emitted by a gas turbine engine while also detuning the exhaust flow to reduce unwanted infrasound. The gas turbine engine test cell includes a test cell building, a thrust frame for mounting the gas turbine engine, and the turbine detuner disposed downstream of the thrust frame for extracting energy from the exhaust gases of the gas turbine engine when in operation. The turbine detuner has an inlet for receiving the exhaust gases, a kinetic energy recovery mechanism (e.g. stator and rotor) for converting the kinetic energy of the exhaust gases into rotary power, and an outlet through which de-energized exhaust gases are emitted after being de-energized by the kinetic energy recovery mechanism. By eliminating the augmentor, the test cell is more compact. The turbine detuner not only reduces infrasound but also recovers otherwise wasted energy. | 06-16-2011 |
20110138773 | DEVICE FOR MOUNTING A FLAME-HOLDER ARM ON AN AFTERBURNER CASING - A flame-holder device for an afterburner of a turbofan, the turbofan including a first inner annular casing and a second inner annular casing, these defining a passage for a primary flow, and an outer annular casing defining with the first inner annular casing a passage for a secondary flow, including at least one arm-support made of metallic material configured to be fixed to the outer casing by an upper plate and at least one flame-holder arm of a one-piece structure of composite material including two mutually attached walls arranged to define a groove whose profile is approximately a V. The walls support, on their upper part situated in the secondary flow, a fixing mechanism for fixing to the arm-support. The upper parts are approximately planar and are positioned facing each other. | 06-16-2011 |
20110146229 | Integration of a Surface Heat Exchanger to the Wall of an Aerodynamic Flowpath by a Structure of Reinforcement Rods - Setting in place of one or several coolers in the wall of a secondary flow of a bypass turbomachine. The wall extends from an intermediate casing toward a leading edge of a separator nose between a primary flow and the secondary flow. The wall includes a series of support arms attached to an intermediate casing, distributed over the perimeter of the wall and directed upstream. A series of surface air-oil heat exchangers forming wall segments are arranged end-to-end on the support arms, so as to form an annular wall. A shroud having a leading edge is arranged and fixed in the area of the upstream edges of the heat exchangers, so as to complete the wall. The support arms include hydraulic connectors connected to one another on each arm, adapted to cooperate with corresponding connectors in the area of the heat exchangers and in the area of the intermediate casing. | 06-23-2011 |
20110154805 | POWER AUGMENTATION SYSTEM FOR AN ENGINE POWERED AIR VEHICLE - One embodiment of the present invention is a unique augmented gas turbine engine propulsion system. Another embodiment is a gas turbine engine power augmentation system. Yet another embodiment is a system for augmenting power in an engine powered air vehicle. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for fluid driven actuation systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith. | 06-30-2011 |
20110173949 | JET ENGINE NACELLE HAVING DAMPERS FOR HALF-SHELLS - The present invention relates to a jet engine nacelle ( | 07-21-2011 |
20110192134 | AIR INTAKE STRUCTURE FOR A TURBINE ENGINE NACELLE - The invention relates to an air intake structure ( | 08-11-2011 |
20110203254 | GAS EJECTION CONE FOR AN AIRCRAFT TURBOJET EQUIPPED WITH A DEVICE FOR GENERATING TURBULENCE IN A PRIMARY FLOW LIMITING JET NOISE - A gas ejection cone for an aircraft turbojet, the cone including a hollow main body defining, on the outside, a radially inner skin of an annular primary flow channel, and a device generating turbulence in the primary flow limiting jet noise, mounted so as to move on the main body so as to be able to be displaced from an extracted position in which the device projects toward downstream in relation to a downstream end of the hollow main body, and a retracted position in which the device is retracted into the hollow main body, and vice versa. Further, the device includes a cylindrical support body having an axis parallel to an axis of the ejection cone, and at least one fin supported by the body. | 08-25-2011 |
20110203255 | FLEXIBLE ABUTMENT LINKS FOR ATTACHING A PART MADE OF CMC - An after-body assembly for an aeroengine, the assembly including, in an axial direction, an annular part made of metallic material secured to the aeroengine and an after-body part made of ceramic matrix composite material presenting the form of a body of revolution at least in its upstream portion, the after-body part being mounted on the annular part by resiliently flexible fastener tabs, each fastener tab having a first end fastened to the annular part and a second end fastened to the upstream portion of the after-body part, wherein each fastener tab includes an axial abutment element extending radially from the second end of the tab, the axial abutment element facing the first end and a radial abutment element at the second end of the tab, the radial abutment element overlying the first end in a radial direction. | 08-25-2011 |
20110209458 | Aircraft gas turbine engine - The invention refers to an aircraft gas turbine engine including a core engine | 09-01-2011 |
20120000179 | REAR SECTION OF AIRCRAFT NACELLE AND NACELLE EQUIPPED WITH SUCH REAR SECTION - The invention relates to a rear section ( | 01-05-2012 |
20120005999 | VARIABLE AREA FAN NOZZLE AND THRUST REVERSER - An example nozzle for use in a gas turbine engine includes a nozzle door having a first end, a second end opposed from the first end, and a first pivot on the door between the first and second ends. A linkage connects to the nozzle door and to an actuator. The actuator is selectively operative to move the linkage about a second pivot to in turn move the nozzle door about the first pivot between a plurality of positions, such as stowed, intermediate, and thrust reverse positions to influence bypass airflow through a fan bypass passage. The fan bypass passage has a radially outward side defined by a nacelle in at least one position. At least a portion of the nozzle door is radially outward of the nacelle and includes a lip extending from the nozzle door between the first pivot and one of the first end and second end. | 01-12-2012 |
20120060465 | DEVICE FOR SPRAYING A FLUID USING THE AIR BLAST EFFECT - A device making it possible to better spray the fluid that one wishes to spray by using a portion of the air blast which is intended to diffuse it. The device has a low pressure airflow generator, supplied with ambient air, capable of creating a high-flow-rate compressed air blast in an aerodynamic stream, and a pre-mixing chamber, supplied with compressed air by tapping a portion of the flow at the outlet of the stream using nozzles, provided with aerodynamic swirlers of which the effect are known, supplement the spraying of the fluid carried out by pressurized ejectors. | 03-15-2012 |
20120060466 | AIR-OIL HEAT EXCHANGER - A heat exchange system for use in fluid operated equipment to provide air and working fluid heat exchanges to cool the working fluid in airstreams on a stream side of a wall. An actuator is mounted to be substantially located on a side of the wall opposite the stream side thereof having a positionable motion effector. A heat exchanger core having a plurality of passageway structures therein to enable providing the working fluid to, and removal therefrom. The heat exchanger core is mounted on the motion effector so as to be extendable and retractable thereby through the opening for selected distances into that region to be occupied by the airstreams. | 03-15-2012 |
20120096831 | GAS TURBINE ENGINE WITH VARIABLE AREA FAN NOZZLE - A nacelle assembly for a bypass gas turbine engine includes a variable area fan nozzle having a first fan nacelle section and a second fan nacelle section. The variable area fan nozzle is in communication with a fan bypass flow path, the first fan nacelle section defines an intermittent trailing edge which defines a multiple of ports and the second fan nacelle section defines a multiple of doors, each of the multiple of doors match each of the multiple of ports such that a fan nacelle trailing edge is continuous when the second fan nacelle section is selectively translated to a closed position relative to the first fan nacelle section. | 04-26-2012 |
20120110979 | FLUTTER SENSING SYSTEM FOR A GAS TURBINE ENGINE - An exemplary gas turbine engine assembly includes a fan casing within a nacelle, a variable area fan nozzle. A controller is operable to move the variable area fan nozzle to influence a discharge airflow area associated with the variable area fan nozzle in response to an airfoil flutter condition. A gear train reduces a rotational speed of a fan in the gas turbine engine relative to another portion of the gas turbine engine. | 05-10-2012 |
20120110980 | VARIABLE AREA FAN NOZZLE FAN FLUTTER MANAGEMENT SYSTEM - A system and method of controlling a fan blade flutter characteristic of a gas turbine engine includes adjusting a variable area fan nozzle in response to a neural network. | 05-10-2012 |
20120117939 | TURBOSHAFT ENGINE WITH REDUCED NOISE EMISSION FOR AIRCRAFT - According to the invention, a plurality of embossments ( | 05-17-2012 |
20120167550 | THRUST AUGMENTED GAS TURBINE ENGINE - One embodiment of the present invention is a unique thrust augmented gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for thrust augmented gas turbine engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. | 07-05-2012 |
20120167551 | ELECTRIC TURBINE BYPASS FAN AND COMPRESSOR FOR HYBRID PROPULSION - An electric turbine compressor fan for hybrid propulsion wherein the compressor contains one or more rotor stages (compressor and diffuser), each being driven by one or more electric ring motors, such that the compressor rotor stages are designed and tuned more precisely to the compression ratio to be attained within the turbine design operating characeristics, thrust requirements and flight envelope. | 07-05-2012 |
20120192544 | HEATED BOOSTER SPLITTER PLENUM - A splitter apparatus for a gas turbine engine includes: a splitter including: an annular outer wall which defines a convex-curved leading edge at a forward end thereof; an annular floorplate positioned radially inboard of the outer wall; and an annular first bulkhead spanning between the outer wall and the floorplate. The outer wall, the floorplate, and the bulkhead collectively define an annular splitter plenum positioned adjacent the leading edge of the outer wall. At least one exhaust passage formed in the outer wall extends past the floorplate and communicates with the exterior of the splitter. At least one jumper tube assembly passes through the first bulkhead, each configured to pass air flow from the exterior of the splitter into the plenum. | 08-02-2012 |
20120222396 | JET ENGINE DEVICE WITH A BYPASS DUCT - A jet engine has a bypass duct limited by an inner wall and an outer wall and inside which a fluid flows. Between the inner and outer walls of the bypass duct a support unit is provided that includes strut-like support elements connected at opposite ends to the inner and outer walls, respectively. Central longitudinal planes of the support elements describe in the areas of the support elements facing the inner wall a positive acute angle with an engine axis, and in the areas of the support elements facing the outer wall a negative acute angle with the engine axis. Flow cross-sections are each enlarged in the area between the side surfaces of the support elements each describing an acute angle with the walls, starting from the areas facing the fluid flow in the direction of the areas of the support elements facing away from the fluid flow. | 09-06-2012 |
20120233980 | VARIABLE CYCLE GAS TURBINE ENGINE - One embodiment of the present invention is a unique variable cycle gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for variable cycle gas turbine engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. | 09-20-2012 |
20120247083 | POROUS CORE COWLING FOR A TURBOJET ENGINE - A core cowling for a bypass turbojet engine consisting of an internal wall ( | 10-04-2012 |
20120260623 | COUPLING SHAFT FOR GAS TURBINE FAN DRIVE GEAR SYSTEM - A coupling shaft assembly for a gas turbine engine includes a forward coupling shaft section with a forward interface spline and a forward mid shaft interface spline and an aft coupling shaft section includes an aft mid shaft interface spline and an aft interface spline, the aft mid shaft interface spline engageable with the forward mid shaft interface spline. | 10-18-2012 |
20130000273 | GAS-DRIVEN PROPULSOR WITH TIP TURBINE FAN - An engine comprises a gas generator having an exhaust plenum and a propulsor comprising a propulsion fan coaxially mounted within an annular turbine. The annular turbine comprises a turbine duct and a plurality of turbine rotor blades rotationally coupled to the propulsion fan. A plurality of hollow struts extend axially and radially from the gas generator to the annular turbine. The hollow struts comprise flow ducts connecting the exhaust plenum to the turbine duct. | 01-03-2013 |
20130014488 | EFFICIENT, LOW PRESSURE RATIO PROPULSOR FOR GAS TURBINE ENGINES - A gas turbine engine includes a spool, a turbine coupled to drive the spool, and a propulsor that is coupled to be driven by the turbine through the spool. A gear assembly is coupled between the propulsor and the spool such that rotation of the turbine drives the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extends from the hub. The row includes no more than 20 of the propulsor blades. | 01-17-2013 |
20130025257 | THREE SPOOL ENGINE BEARING CONFIGURATION - A disclosed gas turbine engine includes a core section having a low spool, intermediate spool and a high spool that rotate about a common axis. The intermediate spool is supported at an aft position by an inter-shaft bearing arrangement on the low spool. The low spool is supported for rotation in an aft position by an aft roller bearing supported on a turbine exhaust case of the gas turbine engine. The high spool is supported by a high spool aft roller bearing disposed within a high spool bearing compartment. The high spool bearing compartment is positioned within a radial space between the combustor and the axis. | 01-31-2013 |
20130025258 | GEARED TURBOFAN BEARING ARRANGEMENT - A geared turbofan gas turbine engine includes a fan section and a core section. The core section includes a compressor section, a combustor section and a turbine section. The fan section includes a gearbox and a fan. A low spool includes a low turbine within the turbine section and a forward connection to a gearbox for driving the fan. The low spool is supported for rotation about the axis at a forward most position by a forward roller bearing and at an aft position by a thrust bearing. | 01-31-2013 |
20130047579 | GAS TURBINE ENGINE FAN VARIABLE AREA NOZZLE WITH SWIVABLE INSERT SYSTEM - An example nacelle assembly for a gas turbine engine includes a nacelle defined about an axis and defining a boundary of a fan bypass flow path. A fan variable area nozzle includes a plurality of inserts movably mounted to said nacelle. Each of the multiple of inserts is located at a circumferential position about the nacelle. The multiple of inserts are each independently moveable into the fan bypass flow path relative the nacelle to selectively vary a fan nozzle exit area. | 02-28-2013 |
20130081375 | CONFORMAL INLET APPARATUS FOR A GAS TURBINE ENGINE - An inlet apparatus for a gas turbine engine includes: a fan duct adapted to surround at least one row of rotating fan blades, the fan duct having a circular frontal area, and defining a first inlet plane; and an outer duct surrounding the fan duct, the outer duct including: a first frontal area shape at the first inlet plane which defines, cooperatively with an exterior of the fan duct, at least one lobe through which air can pass; and a second frontal area shape at a second inlet plane located axially downstream from the forward end which is circular, and which defines, cooperatively with an exterior of the fan duct, an annulus through which air can pass. | 04-04-2013 |
20130097996 | RAM AIR FAN INLET HOUSING - A ram air fan inlet housing for containing a ram air fan rotor. The inlet housing includes a flanged surface and an interior surface. The flanged surface is perpendicular to an axis of the inlet housing and defines a flange plane at an axial end of the inlet housing. The interior surface is symmetric about the axis of the inlet housing and includes a flange section, a transition section, an outlet section, a rotor section, and an inlet section. | 04-25-2013 |
20130097997 | RAM AIR FAN DIFFUSER - A diffuser for a ram air fan assembly includes a perforated cone, an inlet ring seal, and an outlet ring seal. The perforated cone has a frustoconical shape symmetrical about an axis of the diffuser. The inlet ring seal is attached to, and axially disposed about, a first end of the perforated cone. The inlet ring seal includes a fan housing connection having a cylindrical shape. The outlet ring seal is attached to, and axially disposed about, a second end of the perforated cone. An average external diameter of the second end is greater than an average external diameter of the first end such that the perforated cone extends away from the inlet ring seal and radially outward from the axis of the diffuser. | 04-25-2013 |
20130104523 | GAS TURBINE ENGINE WITH AUXILIARY FAN | 05-02-2013 |
20130104524 | TURBOFAN WITH GEAR-DRIVEN COMPRESSOR AND FAN-DRIVEN CORE | 05-02-2013 |
20130111873 | AIRCRAFT TURBOJET ENGINE FAN CASING - An aircraft turbojet engine fan casing in the form of a box section is provided by the present disclosure. The box section includes a radially interior wall of which is able to form an internal skin of a cold air flow duct of a nacelle in which said turbojet engine is intended to be mounted, and a radially exterior wall of which is able to form an external skin of said nacelle, the box forming a module forming an entire thickness of the nacelle, and placed between an upstream portion of the nacelle, forming the air intake, and a downstream casing portion, on which a cascade edge of a thrust reverser can be fixed. | 05-09-2013 |
20130152544 | JET ENGINE - A reaction engine is disclosed mainly intended for aviation, but which as described herein can be adapted for an industrial engine, which makes use of spherical chambers and a pressure system in the blades of the rotor-stator unit which permits a perfect adjustment between said blades and the inner face of the stator, preventing pressure losses. While edges likewise articulated in the same way as the blades, execute a labyrinth seal. | 06-20-2013 |
20130167505 | SEAL FOR A VARIABLE AREA FAN NOZZLE - A primary seal assembly for a variable area fan nozzle (VAFN) equipped turbofan engine includes a deformable seal and a seal retainer attached to the seal. The seal includes an inner wall and webs attached to the inner wall and extending transversely there-from. The inner wall and the webs extend circumferentially at least partially around a bypass duct of the turbofan engine. An inner surface of the inner wall interfaces with the VAFN when the VAFN is in the stowed position. The seal is compressed between the VAFN and the seal retainer when the VAFN is in the stowed position. And each of the webs is deformed into a non-planar configuration when the VAFN is in the stowed position. | 07-04-2013 |
20130174533 | MAGNETICALLY COUPLED CONTRA-ROTATING PROPULSION STAGES - A turbomachine comprises a turbine shaft, first and second rotors, first and second propulsion stages, and a magnetic stator. The first rotor is rotationally coupled to the turbine shaft, and coaxially arranged along an axis. The first propulsion stage is rotationally coupled to the first rotor, opposite the turbine shaft. The second rotor is coaxially arranged about the first rotor, and the second propulsion stage is rotationally coupled to second rotor, opposite the turbine shaft and adjacent the first propulsion stage. The magnetic stator is coaxially arranged between the first rotor and the second rotor, forming a magnetic coupling between the and second rotors to drive the second propulsion stage in contra-rotation with respect to the first propulsion stage. | 07-11-2013 |
20130192196 | GAS TURBINE ENGINE WITH HIGH SPEED LOW PRESSURE TURBINE SECTION - A gas turbine engine includes a very high speed low pressure turbine such that a quantity defined by the exit area of the low pressure turbine multiplied by the square of the low pressure turbine rotational speed compared to the same parameters for the high pressure turbine is at a ratio between about 0.5 and about 1.5. | 08-01-2013 |
20130192197 | ANTI-ICING CORE INLET STATOR ASSEMBLY FOR A GAS TURBINE ENGINE - A gas turbine engine is defined wherein the inlet guide vanes leading into a core engine flow path are sized and positioned such that flow paths positioned circumferentially intermediate the vane are sufficiently large that a hydraulic diameter of greater than or equal to about 1.5 is achieved. This will likely reduce the detrimental effect of icing. | 08-01-2013 |
20130192198 | COMPRESSOR FLOWPATH - A core flowpath through a low pressure compressor section of a gas turbine engine includes an outer diameter, which has a slope angle relative to an axis defined by the core flowpath. The slope angle is a slope angle that is operable to prevent flow separation of a fluid passing through the core flowpath. | 08-01-2013 |
20130192199 | GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION - A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. The shaft includes a main shaft and a flex shaft having bellows. The flex shaft is secured to the main shaft at a first end and includes a second end opposite the first end. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supports the shaft relative to the inlet case. The second bearing is arranged radially outward from the flex shaft. | 08-01-2013 |
20130205752 | GAS TURBINE ENGINE WITH SEPARATE CORE AND PROPULSION UNIT - A gas turbine engine includes a propulsion unit mounted to rotate about a first axis, and a core engine mounted to rotate about a second axis, and wherein the first and second axes are non-parallel. A gas turbine engine includes a propulsion unit driven by a free turbine which is adjacent to the propulsion unit and an associated fan, and having a gas generator core engine including a compressor, combustor and turbine section. A method is also disclosed. | 08-15-2013 |
20130219856 | COUNTER-ROTATING LOW PRESSURE TURBINE WITH GEAR SYSTEM MOUNTED TO MID TURBINE FRAME - A gas turbine engine includes a shaft defining an axis of rotation. An outer turbine rotor directly drives the shaft and includes an outer set of blades. An inner turbine rotor has an inner set of blades interspersed with the outer set of blades. The inner turbine rotor is configured to rotate in an opposite direction about the axis of rotation from the outer turbine rotor. A gear system couples the inner turbine rotor to the shaft and is configured to rotate the inner set of blades at a faster speed than the outer set of blades. The gear system is mounted to a mid-turbine frame. | 08-29-2013 |
20130255224 | REVERSE CORE GEAR TURBOFAN - A gas turbine engine has a fan at an axially outer location. The fan rotates about an axis of rotation. The fan delivers air into an outer bypass duct, and across a booster fan positioned radially inwardly of the outer bypass duct. The booster fan delivers air into a radially middle duct, and across a cold turbine into a radially inner core duct being directed into a compressor. From the compressor, air flows axially in a direction back toward the fan through a combustor section, and across an exhaust of the turbine section as directed into the middle duct. A gear reduction drives the fan from a fan drive turbine section. A method is also disclosed. | 10-03-2013 |
20130305686 | COMBINED TURBOJET AND RAMJET ENGINE - A combined engine includes a turbopump including a pump injecting hydrogen into a heater arranged in an outer casing downstream from a central body, and a subsonic turbine driving the pump, which turbine receives partially-expanded hydrogen collected at an outlet from the heater to apply the hydrogen to a supersonic turbine to operate the engine as a turbojet. The hydrogen from the supersonic turbine is collected in tubes inside the central body to be sent to a combustion chamber defined downstream from the central body, while the hydrogen that is partially expanded in the subsonic turbine is sent directly to the combustion chamber via injectors to operate the engine as a ramjet. | 11-21-2013 |
20130333350 | AIRFOIL INCLUDING ADHESIVELY BONDED SHROUD - An airfoil includes an airfoil body that extends between a leading edge and a trailing edge, a suction side and a pressure side, and a first end and a second end. At least one fitting is located on at least one of the first end and the second end. The fitting includes at least one mounting lug. At least one shroud is adhesively bonded to the at least one fitting. | 12-19-2013 |
20140053532 | NACELLE SCOOP INLET - A scoop inlet for use in a gas turbine engine nacelle has a scoop inlet, and a tab extending forwardly of the scoop inlet. The scoop communicates with a downstream flowpath. The tab has at least one opening at a location upstream of the scoop inlet. A nacelle and a gas turbine engine are also disclosed. | 02-27-2014 |
20140075918 | NACELLE LEADING EDGE ELECTRICAL LATCHING SYSTEM - A nacelle structure for a gas turbine engine includes an outer nacelle surrounding a fan section and defining an outer boundary of a bypass flow passage and an inner nacelle surrounding a core engine section and defining an inner boundary of the bypass flow passage. A panel of the inner nacelle is moveable between an open position providing access to the core engine section and a closed position. A lock is supported within the inner nacelle proximate the panel. The lock includes an electric actuator for moving a locking pin between a locked position and an unlocked position. The lock prevents opening and limits deflection of the panel when in the locked position. | 03-20-2014 |
20140083079 | GEARED TURBOFAN PRIMARY AND SECONDARY NOZZLE INTEGRATION GEOMETRY - A disclosed example geared turbofan engine includes a fan section including a plurality of fan blades rotatable about an axis and a core engine section defined about an engine axis. The core engine section includes a primary nozzle including a primary outer diameter at a primary nozzle trailing edge and a primary maximum inner diameter forward of the primary trailing edge. A bypass passage is defined between an inner nacelle surrounding the core engine section and an outer nacelle and includes a secondary nozzle. The secondary nozzle includes an outer diameter at a secondary nozzle trailing edge and a secondary maximum inner diameter forward of the secondary trailing edge. A ratio between the maximum inner diameter of the primary nozzle and an outer diameter at the trailing edge of the primary nozzle and a ratio between the maximum inner diameter of the secondary trailing edge and the outer diameter at the trailing edge of the secondary nozzle are both less than about 0.700. | 03-27-2014 |
20140083080 | Diode Electrical Ground for Fan Blades - A rotor has a rotor body with at least one slot receiving a blade. The blade has an outer surface, at least at some areas, formed of a first material and having an airfoil extending from a dovetail. The dovetail is received in the slot. A diode is in contact with a portion of the dovetail formed of a second material that is more electrically conductive than the first material. The diode is in contact with a rotating element that rotates with the rotor. The rotating element is formed of a third material. The first material is less electrically conductive than the third material. The diode and the rotating element together form a ground path from the portion of the dovetail into the rotor. An engine and a fan blade are also disclosed. | 03-27-2014 |
20140096507 | Bi-Directional Compression Fan Rotor for a Gas Turbine Engine - A fan rotor has a hub, and a plurality of axial flow fan blades extending radially outwardly of the hub. A radial compressor impeller is positioned radially inwardly of the fan blades. The radial compressor impeller has an upstream inlet which extends generally in an axial direction defined by an axis of rotation of the hub. The radial flow compressor impeller has an outlet that extends radially outwardly of the inlet, and into a supply passage for supplying air to a core engine. An engine is also disclosed. | 04-10-2014 |
20140096508 | SYSTEMS AND METHODS INVOLVING MULTIPLE TORQUE PATHS FOR GAS TURBINE ENGINES - A turbofan engine includes a fan, a compressor section, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor, a shaft configured to be driven by the turbine section and coupled to the compressor section through a first torque load path, and a speed reduction mechanism configured to be driven by the shaft through a second torque load path separate from the first load path for rotating the fan. | 04-10-2014 |
20140157754 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine has a fan section, including a fan, a gear arrangement configured to drive the fan section, and a compressor section, including both a first compressor and a second compressor. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio is provided by the combination of the first compressor and the second compressor, with the overall pressure ratio being greater than or equal to about 35. A pressure ratio across the fan section is less than or equal to about 1.50. The fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct. | 06-12-2014 |
20140157755 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine has a fan section, including a fan, a gear arrangement configured to drive the fan section, and a compressor section, including both a first compressor and a second compressor. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio is provided by the combination of the first compressor and the second compressor. The fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct. A bypass ratio which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section is greater than or equal to about 8. | 06-12-2014 |
20140157756 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine has a fan section, including a fan, a gear arrangement configured to drive the fan section, the geared arrangement defining a gear reduction ratio greater than or equal to about 2.6. A compressor section includes both a first compressor and a second compressor. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio is provided by the combination of the first compressor and the second compressor. | 06-12-2014 |
20140157757 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine has a fan section, including a fan, a gear arrangement configured to drive the fan section, and a compressor section, including both a first compressor and a second compressor. A turbine section is configured to drive the compressor section and the gear arrangement. The turbine section includes a fan drive turbine configured to drive the fan section, a pressure ratio across the fan drive turbine being greater than or equal to about 5. An overall pressure ratio is provided by the combination of the first compressor and the second compressor. | 06-12-2014 |
20140165533 | GAS TURBINE ENGINE MOUNTING RING - A casing for an aircraft engine includes an outer ring and an inner hub defining an airflow passage therebetween, the outer ring having an axis defining an axial direction; a plurality of struts arranged in a circumferential array and extending radially from the inner hub to the outer ring to mount the inner hub to the outer ring; wherein the outer ring is defined by a double skin including an axially-extending annular outer skin of sheet metal concentrically surrounding and radially-spaced from an annular inner skin of sheet metal, the outer and inner skins generally parallel to one another, an annular front end ring and an annular rear end ring welded or brazed to the outer and inner skins adjacent respective front and rear edges of the skins to define an annular cavity between them, and the outer ring further comprising a plurality of circumferentially spaced axially-extending ribs interconnecting the outer and inner skins to reinforce the double skins. | 06-19-2014 |
20140165534 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine has a fan section, including a fan, a gear arrangement configured to drive the fan section, and a compressor section, including both a first compressor and a second compressor. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio is provided by the combination of the first compressor and the second compressor, with the overall pressure ratio being greater than or equal to about 35. The fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct. A bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section, is greater than or equal to about 8. | 06-19-2014 |
20140174055 | TURBINE SECTION OF HIGH BYPASS TURBOFAN - A turbofan engine is disclosed and includes a fan and a compressor in communication with the fan section, a combustor, a turbine and a speed reduction mechanism coupled to the fan and rotatable by the turbine. The turbine includes a first turbine section that includes three or more stages and a second turbine section that includes at least two stages. A ratio of airfoils in the first turbine section to a bypass area is less than about 170. | 06-26-2014 |
20140174056 | GAS TURBINE ENGINE WITH LOW STAGE COUNT LOW PRESSURE TURBINE - A gas turbine engine comprises a gear train defined along an axis. A spool along the axis drives the gear train and includes a low stage count low pressure turbine. A fan i s rotatable at a fan speed about the axis and driven by the low pressure turbine through the gear train. The fan speed is less than a speed of the low pressure turbine. A core is surrounded by a core nacelle defined about the axis. A fan nacelle i s mounted at least partially around the core nacelle to define a fan bypass airflow path for a fan bypass airflow. A bypass ratio defined by the fan bypass passage airflow divided by airflow through the core is greater than about ten (10). | 06-26-2014 |
20140202133 | GAS TURBINE ENGINE MID TURBINE FRAME WITH FLOW TURNING FEATURES - A gas turbine engine includes first and second stages having a rotational axis. A circumferential array of airfoils is arranged axially between the first stage and the second stage. At least one of the airfoils have a curvature provided equidistantly between pressure and suction sides. The airfoils extend from a leading edge to a trailing edge at a midspan plane along the airfoil. An angle is defined between first and second lines respectively tangent to the intersection of the midspan plane and the curvature at airfoil leading and trailing edges. The angle is equal to or greater than about 10°. | 07-24-2014 |
20140208714 | Thermal Management for Gas Turbine Engine - A gas turbine engine has an inlet duct, which is configured to communicate with an inlet to a compressor. The inlet duct is further configured to communicate air outwardly of an outer casing of the gas turbine engine, and to pass the air along an axial length of the gas turbine engine to cool a component associated with the gas turbine engine. | 07-31-2014 |
20140216003 | GAS TURBINE ENGINE WITH GEARED TURBOFAN AND OIL THERMAL MANAGEMENT SYSTEM - A gas turbine engine includes a fan, a compressor section, a combustion section, and a turbine section. A fan drive gear system is configured for driving the fan at a speed different than the turbine section. A lubricant system includes a lubricant pump delivering lubricant to an outlet line. The outlet line splits into at least a hot line and into a cool line. The hot line is directed primarily to locations in the gas turbine engine that are not intended to receive cooler lubricant. The cool line is directed through one or more heat exchangers at which the lubricant is cooled, and the cool line then is routed to the fan drive gear system. At least one of the one or more heat exchangers is a fuel/oil cooler at which lubricant will be cooled by fuel leading to the combustion section. The fuel/oil cooler is downstream of a point where the outlet line splits into the at least the hot line and the cool line, such that the hot line is not directed through the fuel/oil cooler. A method is also disclosed. | 08-07-2014 |
20140216004 | COMPRESSED AIR BLEED SUPPLY FOR BUFFER SYSTEM - A gas turbine engine includes a fan, a compressor section fluidly connected to the fan, a combustor fluidly connected to the compressor section, a turbine section fluidly connected to the combustor, and a buffer system. The buffer system includes a heat exchanger having a first inlet, a first outlet, a second inlet, and a second outlet. The first outlet is configured to provide a cooled pressurized fluid. The buffer system includes first and second air sources that are selectively fluidly coupled to the first inlet, and a third air source that are fluidly coupled to the second inlet. Multiple fluid-supplied areas are located remotely from one another and are fluidly coupled to the first outlet. The multiple fluid-supplied areas include a bearing compartment. A method and a buffer system are also disclosed. | 08-07-2014 |
20140230403 | GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION - A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A first shaft supports a low pressure compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. A first bearing supports the first shaft relative to the inlet case. A second bearing supports a second shaft relative to the intermediate case. A low pressure compressor hub is mounted to the first shaft. The low pressure compressor hub extends to the low pressure compressor section between the first bearing and the second bearing. | 08-21-2014 |
20140245715 | PRIMARY COWL OF A TURBOFAN COMPRISING A ROTATING RING HAVING MICRO-JETS - The invention relates to a primary cowl for a turbofan comprising a primary body generating a primary stream to be ejected through a primary nozzle, and a secondary body generating a secondary stream to be ejected through a secondary nozzle, the primary cowl being shaped so as to be positioned downstream from the primary body and to define, on the inside of the turbofan, the path followed by the primary stream downstream from the primary nozzle and, on the outside, the path followed by the secondary stream downstream from the secondary nozzle. The primary cowl comprises a coupling to a system for supplying a pressurised gas and at least one perforation for injecting the pressurised gas, through the perforation, into the secondary stream. The primary cowl preferably comprises a ring which has perforations and which is rotated about the axis of rotation of the turbofan. | 09-04-2014 |
20140283500 | THRUST EFFICIENT TURBOFAN ENGINE - A turbofan engine includes a gas generator section for generating a gas stream flow with higher energy per unit mass flow than that contained in the ambient air and a power turbine that converts the gas stream flow into shaft power. The turbofan engine further includes a propulsor section including a fan driven by the power turbine through a geared architecture at a second speed lower than the first speed for generating propulsive thrust as a mass flow rate of air through a bypass flow path. An Engine Unit Thrust Parameter defined as net engine thrust divided by a product of the mass flow rate of air through the bypass flow path, a tip diameter of the fan and the first rotational speed of the power turbine is less than about 0.15 at a take-off condition. | 09-25-2014 |
20140283501 | ELONGATED GEARED TURBOFAN WITH HIGH BYPASS RATIO - A propulsion system includes a fan, a gear, a turbine configured to drive the gear to, in turn, drive the fan. The turbine has an exit point, and a diameter (D | 09-25-2014 |
20150013307 | TURBOFAN JET ENGINE - A turbofan engine includes a core engine, having a high-pressure compressor, a combustion chamber and a high-pressure turbine which are coupled to one another via a high-pressure shaft, at least one fan from which gas is supplied into both a primary flow duct and a secondary flow duct of the turbofan engine, at least one low-pressure turbine arranged behind the core engine, and at least one low-pressure shaft, with each low-pressure shaft coupling a fan to a low-pressure turbine. It has been provided that no low-pressure shaft of the turbofan engine passes through the core engine. | 01-15-2015 |
20150027101 | GEARED GAS TURBINE ENGINE ARCHITECTURE FOR ENHANCED EFFICIENCY - An example gas turbine engine includes, among other things, a geared architecture rotatably coupled to the fan drive shaft, and a high pressure compressor. The gas turbine engine is configured so that a core temperature at an exit of the high-pressure compressor is approximately in a range of about 1150 to about 1350 degrees Fahrenheit at take-off. The gas turbine engine is configured so that an Exhaust Velocity Ratio, defined by a ratio of a fan stream exhaust velocity to a primary stream exhaust velocity, is approximately in a range of about 0.75 to about 0.90. A Bypass Ratio of the engine is greater than about 8.0. | 01-29-2015 |
20150113942 | Phosphor Thermometer with Two Waveguides - A phosphor thermometer is disclosed. The phosphor thermometer may comprise a light source configured to emit an excitation light, and an input waveguide configured to transmit at least a portion of the excitation light from the light source to a temperature sensing end. A phosphor may be located at the temperature sensing end and it may be configured to emit a fluorescence signal upon absorption of at least a portion of the excitation light transmitted by the input waveguide. The phosphor thermometer may further comprise an output waveguide configured to transmit at least a portion of the fluorescence signal from the phosphor to a detector. The detector may determine a fluorescence decay constant from the time dependent decay of the fluorescence signal, and the fluorescence decay constant may be correlated with a temperature. | 04-30-2015 |
20150113943 | GEARED TURBOFAN ENGINE WITH POWER DENSITY RANGE - A gas turbine engine turbine has a high pressure turbine configured to rotate with a high pressure compressor as a high pressure spool in a first direction about a central axis and a low pressure turbine configured to rotate with a low pressure compressor as a low pressure spool in the first direction about the central axis. A power density is greater than or equal to about 1.5 and less than or equal to about 5.5 lbf/cubic inches. A fan is connected to the low pressure spool via a speed changing mechanism and rotates in a second direction opposed to the first direction. | 04-30-2015 |
20150121844 | GEARED TURBOFAN ARRANGEMENT WITH CORE SPLIT POWER RATIO - A gas turbine engine includes a fan section and a compressor section. The compressor section includes both a first compressor section and a second compressor section. A turbine section includes at least one turbine and driving the second compressor section and a fan drive turbine driving at least a gear arrangement to drive the fan section. A power ratio is provided by the combination of the first compressor section and the second compressor section, with the power ratio being provided by a first power input to the first compressor section and a second power input to the second compressor section, the power ratio being equal to, or greater than, about 1.0 and less than, or equal to, about 1.4. | 05-07-2015 |
20150128562 | ENGINE DUCT AND AIRCRAFT ENGINE - A projecting part is formed on an inner circumferential wall surface of a nacelle so that the projecting part projects inward in a diametral direction and extends from a front edge of each of circumferentially-oriented side faces of a bottom pylon. The shape of the projecting part seen from an inner side of the diametral direction is a streamline shape extending in parallel with an engine shaft direction. An apex part at the center of the projecting part is positioned on a rear edge of the bottom pylon. | 05-14-2015 |
20150143794 | Geared Turbofan Engine Gearbox Arrangement - A three-spool turbofan engine ( | 05-28-2015 |
20150329213 | INTEGRAL RAM AIR TURBINE STRUT AND GEARBOX - A ram air turbine is provided that utilizes a one-piece strut. The strut includes an integral gearbox section and an integral drive section. Within the strut, a turbine shaft and a bevel gear engages a driveshaft and a pinion gear, which transfers rotation from the turbine to a generator. The strut may be machined from a single piece of metal, such as aluminum. | 11-19-2015 |
20150345426 | GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE - A gas turbine engine includes a fan rotatable about an axis, a compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor. The turbine section includes a fan drive turbine and a second turbine. The second turbine is disposed forward of the fan drive turbine. The fan drive turbine includes at least three rotors and at least one rotor having a bore radius (R) and a live rim radius (r), and a ratio of r/R is between about 2.00 and about 2.30. A speed change system is driven by the fan drive turbine for rotating the fan about the axis. | 12-03-2015 |
20150345427 | GEARED TURBOFAN ENGINE WITH HIGH COMPRESSOR EXIT TEMPERATURE - A gas turbine engine includes a fan with a plurality of fan blades rotatable about an axis, and a compressor section that includes at least first and second compressor sections. An average exit temperature of the compressor section is between about 1000° F. and about 1500° F. The engine also includes a combustor that is in fluid communication with the compressor section, and a turbine section that is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis. | 12-03-2015 |
20150354502 | ENGINE ARCHITECTURE WITH REVERSE ROTATION INTEGRAL DRIVE AND VANELESS TURBINE - A gas turbine engine comprises a fan, a compressor, a turbine having first stage blades and second stage blades, the first stage blades rotates in a first direction and the second stage blades rotates in a second direction opposite the first direction. The first stage blades are directly adjacent to one another, a drive in operable input connection with the fan and in operable output connection with the first stage blades and the second stage blades, the first stage blades and the second stage blades driving the fan through the drive. | 12-10-2015 |
20160010589 | TWO-PART GAS TURBINE ENGINE | 01-14-2016 |
20160053683 | VARIABLE GEOMETRY INLET SYSTEM - A variable geometry inlet system of an aircraft engine includes an inlet duct. The inlet duct includes at least first and second sections moveable between extended and retracted positions such that the inlet duct defines a variable axial length of an inlet passage for selective flight conditions. The inclusion of acoustic treatment may assist in controlling noise. | 02-25-2016 |
20160115865 | GAS TURBINE ENGINE WITH HIGH SPEED LOW PRESSURE TURBINE SECTION AND BEARING SUPPORT FEATURES - A turbine section of a gas turbine engine according to an example of the present disclosure includes, among other things, a fan drive turbine section, and a second turbine section. The fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed. The second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is faster than the first speed. | 04-28-2016 |
20160123233 | SPLINE RING FOR A FAN DRIVE GEAR FLEXIBLE SUPPORT - A gear assembly support for a gas turbine engine includes a spline ring configured to fit into a case of the gas turbine engine and a flex support. The flex support includes splines for engaging the spline ring and an inner portion attachable to a portion the gear assembly. | 05-05-2016 |
20160123339 | RADIAL FASTENING OF TUBULAR SYNCHRONIZING RINGS - A synchronizing ring assembly for a variable guide vane system is disclosed. The synchronizing ring assembly is disposed about an engine axis and includes a synchronizing ring having a radially outer surface with a clevis bracket mounted to the radially outer surface of the synchronizing ring and defining radially extending openings through the clevis bracket. A plurality of fasteners are configured and adapted to be received within the radially extending openings in the clevis bracket for securing the clevis bracket to the radially outer surface of the synchronizing ring. | 05-05-2016 |
20160131035 | VARIABLE GEOMETRY HEAT EXCHANGER APPARATUS - A heat exchanger apparatus including a surface cooler and a passive automatic retraction and extension system coupled to the surface cooler. The surface cooler having disposed therein one or more fluid flow channels configured for the passage therethrough of a heat transfer fluid to be cooled. The heat transfer fluid in a heat transfer relation on an interior side of said one or more fluid flow channels. The surface cooler including a plurality of fins projecting from an outer surface thereof. The passive automatic retraction and extension system including a thermal actuation component responsive to a change in temperature of at least one of the heat transfer fluid and a cooling fluid flow so as to actuate a change in a geometry of the surface cooler. Further disclosed is an engine including the heat exchanger apparatus. | 05-12-2016 |
20160131043 | TURBINE ENGINE GEARBOX - A gas turbine engine comprises a fan, a compressor, a combustor, and a fan drive turbine rotor. The fan drive turbine drives the fan through a gear reduction. The gear reduction includes at least two double helical gears in meshed engagement. Each of the at least two double helical gears are disposed to rotate about respective axes, and each have a first plurality of gear teeth axially spaced from a second plurality of gear teeth by a spacer. Each of the first plurality of gear teeth has a first end facing the spacer and each of the second plurality of gear teeth has a first end facing the spacer. Each first end of the first plurality of gear teeth is circumferentially offset from each first end of the second plurality of gear teeth. A gear ratio of the gear reduction is greater than about 2.3:1. A method is also disclosed. | 05-12-2016 |
20160131084 | GEARED TURBOFAN ARRANGEMENT WITH CORE SPLIT POWER RATIO - A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a compressor section including at least a first compressor section and a second compressor section. A power ratio is provided by the combination of the first compressor section and the second compressor section. A method of design a gas turbine engine is also disclosed. | 05-12-2016 |
20160131151 | GAS TURBINE ENGINE AIRFOIL - An airfoil arrangement of a turbine engine according to an example of the present disclosure includes adjacent airfoils including pressure and suction sides extending in a radial direction from a 0% span position to a 100% span position. The airfoils have a relationship between a gap/chord ratio and span position that defines a curve with a gap/chord ratio having a portion with a negative slope. | 05-12-2016 |
20160186690 | NON-AXISYMMETRIC FAN FLOW PATH - A gas turbine engine propulsion system and method of assembling such is disclosed. The gas turbine engine propulsion system comprises a gas turbine engine that includes a fan flow path. The fan flow path may extend from the fan inlet to the rear exhaust outlet of the bypass flow path. A portion of the fan flow path, proximal to the fan, is non-axisymmetric. The non-axisymmetric portion may be upstream or downstream of the fan. | 06-30-2016 |
20160377027 | EFFICIENT, LOW PRESSURE RATIO PROPULSOR FOR GAS TURBINE ENGINES - A gas turbine engine includes a gear assembly, a bypass flow passage, and a core flow passage. The bypass flow passage includes an inlet. A fan is arranged at the inlet of the bypass flow passage. A first shaft and a second shaft are mounted for rotation about an engine central longitudinal axis. A first turbine is coupled with the first shaft such that rotation of the first turbine is configured to drive the fan, through the first shaft and gear assembly, at a lower speed than the first shaft. The fan includes a hub and a row of fan blades that extend from the hub. The row includes 12 (N) of the fan blades, a solidity value (R) that is from 1.0 to 1.2, and a ratio of N/R that is from 10.0 to 12.0. | 12-29-2016 |
20160377165 | LUBRICANT DELIVERY SYSTEM FOR PLANETARY FAN DRIVE GEAR SYSTEM - A gear system for a turbofan engine assembly includes a sun gear rotatable about an engine centerline, a non-rotatable ring gear, a rotating carrier that drives a fan, and a plurality of planet gears intermeshed between the sun gear and the ring gear. Each of the plurality of planet gears supported on rolling element bearings fit into the carrier. Each of the plurality of planet gears includes an inner cavity and a lubricant passage directed at the rolling element bearings. The carrier includes an outer scoop that receives lubricant from an outer fixed lubricant jet and feeds lubricant into the inner cavity and through the lubricant passage to spray lubricant on to the rolling element bearings. A geared turbofan engine assembly is also disclosed. | 12-29-2016 |
20160377166 | JOURNAL BEARING FOR ROTATING GEAR CARRIER - A gear system for a geared turbofan engine includes a sun gear, a planet gear supported in a carrier and engaged to the sun gear, a forward journal bearing and an aft journal bearing both supporting the planet gear. The carrier includes a forward wall supporting the forward journal bearing and an aft wall supporting the aft journal bearing. A ring gear is engaged to the planet gear. A geared turbofan engine is also disclosed. | 12-29-2016 |
20160377167 | ROLLING ELEMENT CAGE FOR GEARED TURBOFAN - A gear system for a turbofan engine assembly includes a sun gear rotatable about an engine centerline, a non-rotatable ring gear, and a rotating carrier that drives a fan. A plurality of planet gears is intermeshed between the sun gear and the ring gear. A rolling element bearing assembly supports rotation of the planet gear on the carrier. The rolling element bearing assembly includes a rolling element between an inner race and an outer race separated by a cage. A first passage for lubricant through the planet gear. A second passage is in communication with the first passage for communicating lubricant through the inner race to an interface between the inner race and the cage. A geared turbofan engine is also disclosed. | 12-29-2016 |
20170234224 | METHOD AND SYSTEM FOR MODULATED TURBINE COOLING AS A FUNCTION OF ENGINE HEALTH | 08-17-2017 |
20180023470 | GEAR TRAIN ARCHITECTURE FOR A MULTI-SPOOL GAS TURBINE ENGINE | 01-25-2018 |
20190145423 | FAN ASSEMBLY OF A GAS TURBINE ENGINE WITH A TIP SHROUD | 05-16-2019 |